During the past decade there has been growing interest within the space industry towards the development of small satellites. Small satellites are typically categorized as picosats (1 kg or less), nanosats (1-10 kg), microsats (10-100 kg) or minisats (100-500 kg) and range in size from softballs to refrigerators. The interest in these satellites is driven by the current constraints of traditional satellites and launch systems. As a result, there has been a significant effort to push satellite technology to smaller sizes and mass, which would enable small satellites to accomplish missions to complement the larger satellites. Examples of such missions include imaging, remote sensing, surveillance, disaster management, and blue force tracking. These missions are achieved by payloads which demand pointing capabilities from the satellites. This requires an attitude control system (ACS) with small actuators that can fit into the volume and mass constraints of small satellites.
Traditional satellites are typified by budgets in the millions or billions of dollars and schedules on the order of ten years. Failure of traditional satellites is extremely costly hence they tend to utilize space-proven, often outdated technologies, leaving very little room for innovation. An enormous amount of money and effort is expended into the development of redundant systems and the maintenance of outdated techniques and procedures. Thus historically, the development of traditional satellites has been limited to countries with large military and/or commercial budgets.
Small satellites provide an alternative. Improved technologies have allowed small satellites to accomplish many of the tasks of the larger predecessors and at a fraction of the cost and time required of a traditional space satellite [1]. As a result, risk aversion is reduced and small satellite developers are more willing to explore new, unproven technologies that may result in total mission cost reduction and/or increased functionality of the satellite. It should be noted that the willingness of small satellite developers to explore new technologies and innovative designs is not entirely by choice but by the constraints of reduced resources; i.e., “necessity is the mother of invention.” This virtuous circle also has additional benefits in that small satellite developers are now leveraging technologies developed by other non-aerospace industries and are thus helping to underwrite the development cost, i.e., by utilizing commercial off-the-shelf (COTS) items, further cost reductions are being experienced in the space industry.
The advent of small satellites has made space much more accessible and as a result, there has been a large number of proposed (and in development) educational, research, and entrepreneurial satellite efforts [2], [3], [4], [5]. To further expedite the small satellite developmental cycle a standardized platform, known as the CubeSat, was developed [6]. CubeSats are in the picosat class and thus are limited to a mass of 1.3 kg and dimensions of a 10×10×10 cm cube. Multiple CubeSats can also be combined to form nanosats with mass constraints up to 4 kg and dimensions of up to 10×10×30 cm.
A review of the existing and proposed missions illustrates a deficiency in three-axis attitude control of the pico and nano-class satellites [7]. To date, these satellite classes typically use coarse three axis attitude control, or passive control that stabilizes the satellite but fails to provide pointing accuracy. Nadir pointing, when one end or face of the satellite is facing the earth, has been accomplished using either passive means—gravity gradient (e.g. ICE CUBE 1&2, SMART SAT) or magnetic torquers (e.g., DTU SAT, AAU SAT, CUTE 1.7). In cases which required three-axis attitude control, a combination of magnetic torquers (two axes) and a reaction wheel (one axis) have been employed. While these systems have provided some form of attitude control for their experimental mission, the real utility of pico- and nano-satellites cannot be fully realized with these systems.
Control of the angular motion of the spacecraft is inherently management of the vehicle's angular momentum. The angular momentum of the spacecraft can be manipulated by (i) application of an external torque or (ii) by redistribution of the angular momentum within the spacecraft. Both of these effects are captured by the following equation, wherein the latter case is represented by τext=0.Στext=J{dot over (ω)}+ ωJω
The two major components of the ACS are the actuator and the control algorithm. A brief description of the most common attitude actuation systems found in spacecraft applications is presented below. Various types of actuators include the reaction wheel, magnetic rods, torque coils, thrusters, momentum wheels and control moment gyroscope. There are four different types of actuators typically utilized in satellite attitude control. For small satellite (smallsat) mission applications, each of these actuator types has advantages and disadvantages which are discussed in the following paragraphs.
Reaction control devices use spatially distributed linear thrusters to generate external torque. Typically, the linear thrusters are either chemically based or electrically based. The chemically based thrusters utilize a chemical reaction which accelerates a propellant and expels it from the spacecraft. A variant of this system uses pressurized monopropellant to accomplish the same effect. Both of these systems have limited operational life since the momentum transfer is accomplished through expelled particles which must be stored onboard the spacecraft and once expelled are non-reusable. These systems require propellant storage, plumbing, and valves for operations which makes them impractical for pico-/nano-satellite applications. Additionally, these systems require multiple sets of actuators for rapid and precision attitude control.
The electrically based thrusters expel ionized particles accelerated by an electric field which make them more fuel-efficient than their chemical counterparts. Ion thrusters produce less thrust than chemical thrusters and have been implemented as a subset of multiple reaction control attitude actuators. In their current state of development, ion thrusters are impractical for three-axis attitude control due to volume and mass constraints. Furthermore, they may be incapable of performing rapid slew maneuvers due to their low thrust output.
Magnetic actuators usually include either permanent magnets (passive) or coils through which current flows in order to generate magnetic fields (i.e. local B≠0). Magnet torquers generate external torques on the spacecraft by the interaction of the onboard field with the Earth's magnetic field. Since the external torque is a resultant of a vector cross product between the fields, full three axis control is not always possible. Magnet torquers are utilized by CubeSats for de-tumbling where as permanent magnets are typically used for nadir pointing missions. Magnet torquers are ideal for CubeSats since they contain no moving parts, require reasonable power and consume relatively low mass and volume. However, they suffer from (i) spatial dependence (i.e., require knowledge of the local magnetic field), (ii) relatively low accuracy, and (iii) singular directions potential attitudes for which magnetic coils become ineffective).
Momentum exchange devices include flywheels that are either spinning at high speeds (momentum wheels—MW), accelerated (reaction wheel—RW), or moderately spinning wheels that are gimbaled (control moment gyroscopes—CMG). All three systems are susceptible to momentum saturation (e.g., from external disturbances) and require some capabilities for momentum dumping. Typically, the momentum dumping is accomplished with either magnet torquers or reaction control devices.
Momentum wheels differ from reaction wheels in that RWs have a zero nominal operating speed. Reaction wheels as the name implies, produce a torque on the spacecraft in response to a torque applied by a motor to a flywheel. These devices are relatively simple in design but require substantial power (i.e., shaft power) due to the direct nature of the application of torque (i.e., the output to input torque ratio is unity) and typically have lower slew rates than CMGs. Combinations of RWs (typically one) and magnetic torquers have been implemented on nano-satellites for 3-axis attitude control.
CMGs rotate the angular momentum along a flywheel axis about a gimbal axis to produce a gyroscopic control torque as shown in FIG. 1. The output torque (gyroscopic torque) is amplified over the input torque required to rotate the gimbal axis (due to the satellite angular velocity) resulting in the well known torque amplification factor which allows for higher slew rates. This property of torque amplification as well as the fact that CMGs require minimal shaft power, permits the CMG to have a much higher torque per unit mass and unit power ratio than RWs.
More specifically, the CMG is a mechanism that produces torque by a combination of two motions—spinning a flywheel about an axis referred to as the flywheel axis and the rotation of the spinning flywheel about an axis perpendicular to flywheel axis referred to as the gimbal axis. The two main components of a gyroscope are the flywheel and the gimbal. The flywheel is a spinning rotor with inertia sufficient to provide the desired angular momentum; the gimbal is a pivot about which the flywheel assembly can be rotated. The magnitude of the gyroscopic torque produced is directly proportional to the inertia of the flywheel, the angular speed of the flywheel and the rate of rotation of the gimbal. In a CMG, the inertia of the flywheel and the speed of the flywheel are constant, and the torque output is controlled by changing the rotation rate of the gimbal. The direction of the torque produced is perpendicular to both the flywheel and the gimbal axes per the right hand rule. This torque acts on the satellite structure to change its attitude. A combination of gyroscopes is used to produce a net torque in the desired direction and magnitude. There are various combinations of gyroscopes that can be used depending upon the mission requirements (box configuration, inline configuration, roof top configuration, pyramidal configuration).
Apart from the gyroscopic torque produced by the CMG, there are other torques that arise from the motion of the flywheel and gimbal that contribute to the dynamics of the satellite:                Reaction torque due to friction in the flywheel bearings.        Reaction torque due to the acceleration of the gimbal; this torque depends on the angular acceleration and the inertia of the gimbal.        Reaction torque due to the friction of the gimbal bearings and slip ring.        
The motion to the flywheel and gimbal is provided by flywheel and gimbal motors. There are feedback devices (e.g., encoders and Hall-effect sensors) for sensing the angular speed and position. A slip ring is provided for continuous power supply to the flywheel motor for endless rotation of the gimbal. All these hardware are assembled together with structural components.
The CMG shown in FIG. 1 is in its basic form and called the single gimbal control moment gyroscope (SGCMG). The torque output of this CMG is in a unique direction for every orientation of the gimbal and flywheel axis. The torque span of this type of CMG lies in a plane (for 360° rotation of the gimbal axis). The SGCMG is popular and widely used for its simplicity in mechanical construction and relatively simpler control logic.
The second type of CMG is the double gimbal control moment gyroscope (DGCMG). In this type there are two gimbals about which the flywheel assembly can rotate. The output torque direction of this CMG is determined by the angular positions of both the gimbals and since these gimbals are in two different orthogonal planes, the torque output is in 3D space and not confined to a plane as in a SGCMG. One of the drawbacks of this type is the phenomenon of gimbal lock which occurs when the flywheel and gimbal axes align. In this situation the CMG cannot produce any torque. The mechanical construction of the DGCMG is more complex.
Another type of CMG is the variable speed control moment gyroscope (VSCMG). This CMG controls the acceleration of the flywheel to produce torque in addition to the gyroscopic torque produced by gimbal movement. The output torque direction of this CMG is determined by the acceleration of the flywheel and the orientation of the gimbal. The torque span hence lies in 3D space. Two different control algorithms—one for the flywheel and the other for the gimbal needs to integrated for the functioning of the VSCMG.
Therefore, there exists a need for a CMG capable of rapid retargeting (e.g., high slew rates) and attitude control of small satellites (e.g., pico and nano-satellites) using a compact actuation system. Moreover, there is a need for a CMG meeting various size and performance constraints, such as mass, power, and volume constraints for these small satellites.